Conventional systems for attitude determination of aerospace vehicles include star trackers and gyroscopes. Using a star tracker alone to determine attitude has disadvantages. If the spacecraft is undergoing high slew rates, i.e. is moving too fast or tumbling, the star camera field of view (FOV) may be changing too fast, and the star tracker camera may not be able to focus on the FOV. The position of stars and the brightness of stars is therefore blurred and cannot be resolved by the star camera sufficiently to be properly compared to a star catalog of known star patterns. Also, if the star camera system has no prior knowledge of the attitude of the aerospace vehicle, it must rely on use of a “lost-in-space” algorithm, where the star identification process is computationally more intensive because of the lack of base information regarding prior aerospace vehicle attitude. Also, if the Sun, Earth, Moon or other bright stars enter the star camera field-of-view during the imaging process, the camera field-of-view is occluded, i.e. the star pattern is obscured because the intensity of the stars in the star pattern is less than the intensity of the Sun, Earth or Moon. This results in poor resolution and consequent inability to properly match the star pattern in the star camera field-of-view with star patterns in the star catalog.
The use of only a gyroscope for attitude determination has disadvantages as well. Inherent errors associated with gyroscopes as known by those skilled in the art include gyroscope “drift”, bias and scale factor errors, which affect the accuracy of the gyroscope output.
Additionally, known systems use a Charge Coupled Device (CCD) imager for the star camera, and conventional gyroscopes. Their size, weight, control electronics, and power requirements of CCDs prohibit their use in small satellite applications. CCDs are also susceptible to radiation damage and conventional gyroscopes suffer from the further disadvantages of size and weight, sensitivity to vibrations, and susceptibility to radiation.
Despite these disadvantages, the combination of a CCD star tracker camera attitude determination system and a separate conventional gyroscope attitude determination system is a widely accepted method of solving the attitude determination problem. The combination of CCD star camera systems and conventional gyroscope systems decreases the effects of image blurring and lack of resolution when the vehicle's attitude rate exceeds the camera's optical tracking limit. It also lessens the effects of occlusion of the star field image caused by the presence of the Sun, Earth or Moon in the camera's field-of-view. This combination accomplishes these goals by relying on the gyroscope output to determine the aerospace vehicle attitude when the star camera is occluded or if the aerospace vehicle is moving too fast or is tumbling. Conversely, the attitude derived from the star camera output may be used to calibrate the gyroscope system errors.
Traditionally both the star camera and gyroscope systems obtain attitude information separately, and each system outputs its attitude information to a spacecraft's flight computer where application-specific software must be created to resolve the two separate attitude inputs. This is a cumbersome, highly power and computationally intensive and expensive process that is prone to error. Also, the star camera system and the gyroscope system are separate from each other structurally, adding weight to the overall aerospace vehicle attitude determination system.
Generally the star camera system and the gyroscope system are each developed and manufactured by different companies, with each system having different hardware and software. When loaded on board an aerospace vehicle, the separate star camera and gyroscope systems require that the spacecraft's flight computer fuse or integrate the data generated by the two systems “externally” or outside of both of these systems, using custom mission-specific software. The software designer is thus required to have detailed knowledge and understanding of each systems' behavior, down to the intricacies of how changes in the systems may occur over time, by reason of temperature variations, or as a consequence of other environmental factors. Integrating each system's outputs in software creates unnecessary design risk because the engineers performing the integration are not the same engineers who designed and best understand each system's hardware. Furthermore, third party integration and lack of insight into the design of each system leads to less than optimum performance. Also, this “external” fusion of gyroscopes and star trackers requires large mass and high power, which prohibits applications in smaller vehicles such as nanosatellites. Spacecraft of less than ten kilograms cannot use traditional gyroscopes and star tracker cameras at all. Moreover, the cost of the separate systems, together with the cost of integrating the two systems externally, is quite expensive.